This invention relates to the prevention of electrical charge buildup on a spacecraft.
Many current spacecraft include a body and large solar cell arrays, which are folded against the body for launch and thereafter deployed from the body as "wings". The solar cell arrays produce electrical power for the spacecraft. They are typically quite large in extent for spacecraft having high power consumption, such as geosynchronous communications satellites.
When a spacecraft is operated in space, imbalances in the collection and emission of electrical charge from the ambient environment can result in absolute charging between the spacecraft and its ambient plasma environment. Due to differences in the properties of exterior surface materials, the electrical charging can also cause differential charging between different parts of the spacecraft. If the electrical charges become sufficiently great, the resulting high voltages and electrical arcing may adversely affect the performance of electrical equipment on the spacecraft, degrade performance of the solar cell array, or even result in damage to the spacecraft structure.
Absolute electrical charging is a by-product of the interaction of a spacecraft with the varying flux of electrons and ions in the ambient space environment. A neutral surface in space collects a higher flux of electrons than positive ions. A spacecraft charges negatively in response to the net collection of negative charge, which enhances the collection of more ions and suppresses the collection of electrons. The negative charging continues until the net collection of ions and electrons reaches an equilibrium. A large geosynchronous spacecraft in eclipse can charge to 5,000-20,000 volts under disturbed environment conditions. Photoemission from sunlit surfaces returns a large flux of electrons to the ambient environment, offsetting the collection of electrons by surfaces that are in shadow. The electrically conductive portion of the shaded surface of the spacecraft tends to collect electrons from the plasma environment, and the electrically conductive portion of the illuminated surface tends to photoemit electrons into the plasma environment. During calm ambient environment conditions, photoemission from the conductive surfaces tends to emit more electrons than collected from the shadowed conductive surfaces, which causes a geosynchronous satellite to acquire a small positive charge in the range of a few volts. During disturbed, active ambient environment conditions, the electron collection from shadowed surfaces can overwhelm photoemission, leading to charging of the spacecraft to potentials in the kilovolt range.
Non-conductive materials in shadow collect electrons from the environment and charge negatively relative to the underlying structure, because the collected charge cannot flow to sunlit surfaces where it is photoemitted. Non-conductive sunlit materials photoemit large numbers of electrons and charge positively relative to the underlying structure because electrons cannot flow from the structure to the surface of the material. In both cases, differential voltages can result that lead to electrical arcs. Differential electrical charging between different parts of the spacecraft is controlled by providing the external surfaces of the spacecraft with sufficient electrical conductivity to equalize the charges over the surface. This electrical conductivity is imparted through the use of coatings on the surface materials of the spacecraft or the use of bulk additives within the surface materials. In both cases, the materials must acquire sufficient electrical conductivity to dissipate static charges.
Both passive and active approaches have been developed to aid in maintaining a neutral charge between the spacecraft and the plasma environment. Because the absolute charging phenomenon is related to the relative amounts of surface area that are shaded from the sun and are illuminated by the sun, the passive approaches must take into account the photoelectric behavior of the large-area solar cell arrays. The satellite is oriented so that the solar cells of the arrays always face the sun, so that there is a "front" side of the solar cell array facing toward the sun, and a "back" side facing away from the sun. The back side of the solar cell array is electrically conductive and accepts electrons from the plasma environment. The front side of the solar cell array includes a protective cover over the solar cells made of glass and other electrically nonconductive materials. Absent some further treatment for the front side of the solar cell array, the electrons collected on the back side of the array cannot flow to the surface of the electrically nonconductive covers on the front side, which would then photoemit the electrons. As a result, net absolute charging of the spacecraft occurs along with differential charging between the solar cell covers and the rest of the solar cell array.
To alleviate this charging, it has been known to apply a thin coating of an electrically conductive, transparent material, such as indium tin oxide (ITO), to the front side of the solar cell array over the cover glass in such a way that an electrical connection is made between the coating and the structure. The ITO-coated cover glass is a passive photoemitter, as was the uncoated glass. By providing a conductive path to the surface with the ITO coating, photoemission from the array front surface can thereby balance the net charge transfer with the plasma environment to prevent charging between the spacecraft and the plasma environment, and differential charging between the covers and the rest of the array. However, the ITO coating is not perfectly transparent and attenuates the sunlight incident upon the solar cell array, reducing its net efficiency. Charging protection is thereby achieved at the cost of reduced electrical production, so that even-larger solar cell arrays are required. As an alternative, active systems for monitoring the plasma environment and for responsively controlling charge accumulation have been proposed, but these tend to add weight to the spacecraft and have not been proved reliable for projected 15-year spacecraft lives.
The present approaches for controlling the buildup of charges and the associated high voltages and arcing are not fully acceptable, and satellite failures and/or degradations continue to occur as a result. Accordingly, there is an need for an improved approach to the control of spacecraft charging relative to the plasma environment. The present invention fulfills this need, and further provides related advantages.